Oil tank filling system

ABSTRACT

An oil tank filling system for filling the oil tank of a gas turbine engine that includes an aft core cowl. The system includes an oil tank that has an oil tank top and an oil tank bottom and is located within a core of the engine, an oil access port located on the aft core cowl, and an oil tank filling pipe that leads from the oil access port to a tank filling port located at or adjacent the oil tank bottom. The system is configured so that oil that is supplied to the oil access port flows to and into the oil tank using gravitational force. A method of filling the oil tank of a gas turbine engine and a gas turbine engine that includes the oil tank filling system.

CROSS-REFERENCE TO RELATED APPLICATIONS

This specification is based upon and claims the benefit of priority fromUnited Kingdom patent application number GB 1807202.5 filed on May 2,2018, the entire contents of which are incorporated herein by reference.

BACKGROUND Field of the Disclosure

The present disclosure relates to an oil tank filling system.

More particularly, the present disclosure provides an oil tank fillingsystem that is useful for a gas turbine engine such as those used topower and propel large passenger aircraft. A method of filling the oiltank of a gas turbine engine and a gas turbine engine that includes theoil tank filling system are also disclosed.

Description of the Related Art

Gas turbine engine oil systems are in general full flow re-circulatorysystems that must provide for adequate lubrication and cooling of allengine bearings, gears and drive splines under all foreseeable operatingconditions.

Each gas turbine engine typically has an oil tank that supplies oil tothe oil system of engine. The oil tank is typically filled through asecurable opening in the tank. For turbo fan engines this opening istypically located on the fan case towards the front of the engine. Forexample for the Rolls-Royce® TRENT® 1000 engine, the oil tank is locatedon the left hand side of the fan case (viewed from the air intake). Anoil filler cap is readily accessed by a maintenance engineer or servicetechnician who mounts a set of steps with an oil can, removes the fillercap and simply pours oil from the oil can into the oil tank.

While that method is effective for those engines and many other engines,it is not suitable, or at least not favourable, for engines wherebroader design reasons require the oil tank to be located within thecore engine. For example, the oil tank may be housed within an enginenacelle and is therefore generally inaccessible to a maintenanceengineer with an oil can or it is unsafe or at least inconvenient to beaccessed by a maintenance engineer. These problems are acute when theoil tank is located below one or more outlet guide vanes (OGV) on thetop of the core engine where room is tight and access is restricted bythe fan case and the outlet guide vanes.

Access is especially restricted within the core engine of large turbofanengines where the large size of the fan requires a large fan case.Furthermore large turbofan engines typically need to be geared, thegears need to remain lubricated and therefore large geared turbofanengines tend to require large oil tanks.

The safety of passengers, crew and also maintenance staff is paramountto the design and operation of gas turbine engines. Furthermore it iscritical to the timely maintenance of a gas turbine engine that its oilsystem can be filled with oil in an efficient manner.

There is therefore a need to provide oil tank filling systems forfilling oil tanks that are located in a part of the engine that is notreadily accessible to maintenance engineers, for example within the coreengine of a turbofan engine.

U.S. Pat. No. 9,194,294 B2 discloses a turbofan gas turbine aircraftengine that has an oil tank in a core compartment and an oil fill tube,sealable with a cap and sequestable under a cover, which is mounted tothe fan case and is fluidly connected to the oil tank via a fluidconduit and a tube. The fluid conduit passes through a dry cavity in theflow exit guide vane. Alternatively oil from the oil fill tube flowsthrough a wet cavity in the flow exit guide vane to reduce piping. Thegas turbine engine requires a maintenance engineer with an oil can tofill the oil tank.

U.S. Pat. No. 8,627,667 B2 discloses a turbofan gas turbine aircraftengine that has a fluid tank structure that is integrated within abypass duct. Oil is added to the tank via a fill tube. The enginerequires a maintenance engineer with an oil can to pour oil in to thefill tube to fill the oil tank.

United States patent application US 2013/0291514 A1 discloses a gasturbine engine that has an oil tank that is arranged near the hotsection of the engine, i.e. in the vicinity of the combustor section andthe turbine section, to address packaging constraints. The oil tank isaxially aligned with the compressor section and filled with oil througha fill tube mounted on an upper portion of the fan case and a tube thatleads to the oil tank.

United States patent application US 2015/075132 A1 discloses anarrangement of an oil tank of an aircraft turbomachine. The oil tank ismounted at a distance from a pump to optimise space dedicated for thetank between an intermediate casing and a nacelle cowling. The tank canbe filled by accessing via an aperture in a panel located on the upperportion of the nacelle cowling. The oil level in the tank can bevisually determined from a gauge on the panel. A technician must reachhigh within a tight space in the engine to access an oil returnconnector.

The present disclosure provides an oil tank filling system thatovercomes at least some of the above problems or at least a usefulalternative to known oil tank filling systems.

SUMMARY

In a first aspect, there is provided an oil tank filling system forfilling an oil tank of a gas turbine engine that includes an aft corecowl, the system comprising: an oil tank that has an oil tank top and anoil tank bottom and is located within a core of the engine; an oilaccess port located on the aft core cowl; and an oil tank filling pipethat leads from the oil access port to a tank filling port located at oradjacent the oil tank bottom; the system being configured so that oilthat is supplied to the oil access port flows to and into the oil tankusing gravitational force.

The system enables an oil tank of a gas turbine engine to be filled whenthat oil tank that is located in a part of the engine that is notreadily accessible e.g. within the core engine, for example moreparticularly within the core engine of a turbofan within the peripheryof the fan case. The aft core cowl is typically readily and safelyaccessible with sufficient space underneath the core cowl to accommodateoil piping. The system also enables the oil tank to be filled by amaintenance engineer whilst avoiding the potential hazards of raisingthe maintenance engineer above the core of the engine.

The oil tank may be shaped to curve around one or both sides of theengine within the core of the engine.

The oil tank may be located between opposing outlet guide vanes.

A sight viewing glass may be provided that shows or indicates the levelof oil in the oil tank.

The sight viewing glass may be located to be visible through an accesspanel on the aft core cowl and may communicate with the oil tank fillingpipe to visually alert when oil is backing up the oil filling tank pipe.

The sight viewing glass may be located to be visible through an accesspanel on the aft core cowl adjacent the oil access port and maycommunicate with the oil tank filling pipe so that the level shownthrough the sight viewing glass reflects that of the oil in the oiltank.

The oil access port and the sight viewing glass may be positioned to beoutboard of a matrix cooler i.e. to take advantage of the cooler'spresence to shield it from radiation from the core, more specificallycasings of the core.

At least one oil sensor may be provided in or adjacent the oil tankcommunicates with an oil sensor controller that provides an alert whenthere is a predetermined volume of oil in the oil tank.

One or more temperature and/or pressure sensors may be providedalongside one or more sections of the oil tank filling pipe and maycommunicate with a controller to provide an alert should the oil tankfilling pipe burst or become blocked.

The oil tank filling system may include an oil pump that may be activeto pump oil through the oil tank filling pipe and the tank filling portand into the oil tank.

The oil tank filling system may include an oil pump controller thatcontrols the operation of the oil pump in the filling of the oil tank.

In a second aspect, there is provided a method of filling an oil tank ofa gas turbine engine that includes an aft core cowl, the methodcomprising the steps of: supplying oil to an oil access port located onthe aft core cowl of the engine; and transporting the oil usinggravitational force from the oil access port through an oil tank fillingpipe that leads to a tank filling port, located at or adjacent an oiltank bottom of an oil tank located within the core of the engine, andinto the oil tank.

In a third aspect, there is provided a gas turbine engine that includesthe oil tank filling system of the first aspect.

The gas turbine engine may be for an aircraft wherein the gas turbineengine comprises an engine core comprising a turbine, a compressor, anda core shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades; and a gearbox that receives an input from the core shaft andoutputs drive to the fan so as to drive the fan at a lower rotationalspeed than the core shaft.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

The term “core” or “core engine” as used herein means the part of a gasturbine engine that houses the compressor(s), combustor(s), turbine(s),and the core shaft(s) that connect the turbine(s) to the compressor(s).The core is typically contained within an engine nacelle.

Throughout this specification and in the claims that follow, unless thecontext requires otherwise, the word “comprise” or variations such as“comprises” and “comprising”, will be understood to imply the inclusionof a stated integer or group of integers but not the exclusion of anyother stated integer or group of integers.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will be described by way of example only with reference tothe accompanying drawings. In the drawings:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of thegas turbine engine;

FIG. 3 is a partially cut-away view of a gearbox for the gas turbineengine;

FIG. 4 is a schematic sectional view of a geared turbofan engine showingan embodiment of the oil tank filling system of the present disclosure.

The following table lists the reference numerals used in the drawingswith the features to which they refer:

Ref no. Feature FIG. A Core airflow 1 B Bypass airflow 1 9 Principalrotational axis (of engine) 1, 2 10 Gas turbine engine 1, 4 11 Core 1, 412 Air intake 1, 4 14 Low pressure compressor 1 15 High pressurecompressor 1 16 Combustion equipment 1 17 High pressure turbine 1 18Bypass exhaust nozzle 1 19 Low pressure turbine 1 20 Core exhaust nozzle1 21 Fan nacelle or fan case 1, 4 22 Bypass duct 1, 4 23 Fan 1, 2, 4 24Stationary supporting structure 2 26 Shaft 1, 2 27 Interconnecting shaft1 28 Sun wheel or sun gear 2, 3 30 Epicyclic gear arrangement 1, 2, 3 32Planet gears 2, 3 34 Planet carrier 2, 3 36 Linkages 2 38 Sun gear 2, 340 Linkages 2 50 Oil tank filling system 4 52 Oil tank 4 55 Oil accessport 4 56 Oil tank filling pipe 4 57 Oil filling level 4 58 Tank fillingport 4 62 Engine nacelle 4 64 Outlet guide vane (OGV) 4 66 Oil tank top4 68 Oil tank bottom 4 70 Lower bifurcation 4 74 Auxiliary gear box 4 76Sight viewing glass 4

DETAILED DESCRIPTION OF THE DISCLOSURE

The present disclosure relates to an oil tank filling system that isuseful, for example, in the gas turbine engine of an aircraft.

Such a gas turbine engine may comprise an engine core comprising aturbine, a combustor, a compressor, and a core shaft connecting theturbine to the compressor. Such a gas turbine engine may comprise a fan(having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed). In some alternative embodiments the core shaft may receivedrive from a turbine without the core shaft also being connected to acompressor.

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

In some embodiments, the or each turbine may be a centrifugal turbine.

In some embodiments, the or each compressor may be a centrifugalcompressor.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm(around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390cm (around 155 inches). The fan diameter may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be inthe range of from 1700 rpm to 2500 rpm, for example in the range of from1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for an engine having a fandiameter in the range of from 320 cm to 380 cm may be in the range offrom 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.3, 0.31,0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in thisparagraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, or 17. The bypass ratio may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The bypass duct may besubstantially annular. The bypass duct may be radially outside the coreengine. The radially outer surface of the bypass duct may be defined bya fan nacelle and/or fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹ s, 105 Nkg⁻¹ s, 100 Nkg⁻¹ s, 95 Nkg⁻¹ s, 90 Nkg⁻¹ s, 85 Nkg⁻¹ s or80 Nkg⁻¹ s. The specific thrust may be in an inclusive range bounded byany two of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15° C. (ambient pressure 101.3 kPa, temperature 30° C.), withthe engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400 K, 1450 K, 1500 K,1550 K, 1600 K or 1650 K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a blisk or a bling. Any suitable method may be used tomanufacture such a blisk or bling. For example, at least a part of thefan blades may be machined from a block and/or at least part of the fanblades may be attached to the hub/disc by welding, such as linearfriction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean cruise conditions of anaircraft to which the gas turbine engine is attached. Such cruiseconditions may be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or engine at the midpoint (in terms of time and/or distance) betweentop of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be the cruise condition. For someaircraft, the cruise conditions may be outside these ranges, for examplebelow Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions at an altitude that is in the range offrom 10000 m to 15000 m, for example in the range of from 10000 m to12000 m, for example in the range of from 10400 m to 11600 m (around38000 ft), for example in the range of from 10500 m to 11500 m, forexample in the range of from 10600 m to 11400 m, for example in therange of from 10700 m (around 35000 ft) to 11300 m, for example in therange of from 10800 m to 11200 m, for example in the range of from 10900m to 11100 m, for example on the order of 11000 m. The cruise conditionsmay correspond to standard atmospheric conditions at any given altitudein these ranges.

Purely by way of example, the cruise conditions may correspond to: aforward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of−55° C.

As used anywhere herein, “cruise” or “cruise conditions” may mean theaerodynamic design point. Such an aerodynamic design point (or ADP) maycorrespond to the conditions (comprising, for example, one or more ofthe Mach Number, environmental conditions and thrust requirement) forwhich the fan is designed to operate. This may mean, for example, theconditions at which the fan (or gas turbine engine) is designed to haveoptimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

An example of a gas turbine engine for which the oil pipe failuredetection system of the present disclosure is useful will now be furtherdescribed with reference to the some of the drawings.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A fan nacelle 21 surrounds the fan case of thegas turbine engine 10 and defines a bypass duct 22 and a bypass exhaustnozzle 18. The bypass airflow B flows through the bypass duct 22. Thefan 23 is attached to and driven by the low pressure turbine 19 via ashaft 26 and an epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The low pressure turbine 19 drivesthe low pressure compressor 14 via shaft 26. The fan 23 generallyprovides the majority of the propulsive thrust. The epicyclic gearbox 30is a reduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32. The planet gears 32are supported for rotation on bearings. The bearings may be of anysuitable kind, such as journal bearings or rolling element bearings.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area. Whilst the described example relates to aturbofan engine, the disclosure may apply, for example, to any type ofgas turbine engine, such as an open rotor (in which the fan stage is notsurrounded by a fan nacelle) or turboprop engine, for example. In somearrangements, the gas turbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

Turning now more specifically to the oil tank filling system of thepresent disclosure that may be used in such a gas turbine engine. Thesystem will be described with reference to FIG. 4.

In broad terms the oil tank filling system 50 comprises an oil tank 52,an oil access port 55, an oil tank filling pipe 56, and a tank fillingport 58.

The oil tank 52 is located within the core 11 of the engine 10, the corebeing housed within an engine nacelle 62. The oil tank will haverestricted access by being within the core of the engine. However eventhat part of the engine nacelle 62 that is adjacent or closest to theoil tank 52 may have restricted access by virtue of the presence ofsurrounding engine structures such as outlet guide vanes 64 or otherequipment and/or components so that an oil inlet formed in that part ofthe engine nacelle 62 would be very difficult and/or potentiallyhazardous to be accessed by a maintenance engineer equipped with asimple oil can. As mentioned above, such access can be especiallyrestricted within the core engine of a large turbofan engine where thelarge size of the fan requires a large fan case and other designconsiderations mean certain equipment is best located on the relevantpart of the core.

The oil tank 52 can take a variety of shapes and configurations for theengine concerned. For example the oil tank may be shaped to curve aroundone or both sides of the engine 10 within the core 11 of the engine.Such an oil tank may be located between outlet guide vanes 64 onopposing parts of the core.

In FIG. 4 the oil tank 52 is shaped to curve around one side of theengine 10 within the core 11 of the engine. The oil tank 52 has an oiltank top 66 located near the top or uppermost part of the core 11 and anoil tank bottom 68 located near the bottom or lowermost part of the core11. Access to the oil tank top 66 of the oil tank 52 is restricted bybeing located within the core 11 of the engine and being locateddirectly beneath an outlet guide vane 64. Similarly access to the oiltank bottom 68 of the oil tank 52 is restricted by being located withinthe core 11 of the engine and being located directly above an outletguide vane 64.

In the oil tank filling system 50 of the present disclosure access tothe oil tank 52, more specifically the oil tank bottom 68, is obtainedvia the tank filling port 58 that is located at or adjacent the oil tankbottom 68, for example within or adjacent a base of the oil tank 52. Inthe embodiment shown in FIG. 4, the tank filling port 58 is formed inthe base of the oil tank 52.

The oil tank filling system includes an oil access port 55 through whichthe oil tank filling system can be charged with oil. The oil access port55, which can take various suitable forms, is located on the aft cowl ofthe engine, which is readily accessible to a maintenance engineer. Thisenables the filling of the oil tank 52 that is located in a part of theengine that is not readily accessible to a maintenance engineer. Inother words, the oil tank filling system of the present disclosureprovides for a generally inaccessible oil tank 52 to be filled remotelyfrom a readily accessible oil access port 55.

The location of the oil access port 55 with respect to the oil tank maydetermine or at least influence the volume of oil that can be added tothe oil tank 52. This is because, in general, when relying on gravityalone to transport oil through the oil tank filling system of thepresent disclosure, the oil of level in the oil tank 52 cannot be higherthan the level of oil in the oil access port 55 when the oil tankfilling pipe 56 is full. For example, in the oil tank filling system 50that is shown in FIG. 4, the oil filling level 57 in the oil tank 52 islevel with the oil access port 55. The system is therefore full of oilalthough the oil tank itself is not completely full of oil.

An air release valve (not shown) may be provided in the oil tank top 66to release air that might accumulate in the oil tank 52 and reduce theavailable volume of oil in the oil tank 52.

A pressure balancing or anti-siphoning pipe (not shown) may connect theoil access port 55 and the oil tank 52 to the balance the air pressurebetween the oil access port 55 and the oil tank 52 (i.e. more especiallywithin the oil tank top 66 above the oil filling level 57).

In some embodiments the oil is provided into the oil access port 55 bysimply pouring oil from an oil can or similar receptacle into the oilaccess port 55. This can be undertaken by a maintenance engineer armedwith such an oil can or similar receptacle but as the oil access port 55is located on the aft core cowl of the engine it will generally benecessary for the maintenance engineer to use a set of steps or beraised on a platform to do this.

In other embodiments the oil is provided into the oil access port 55 viaa suitable, detachably connectable, oil supply apparatus. Such an oilsupply apparatus may include a pump.

In the oil tank filling system of the present disclosure the oil accessport 55 provides access for the oil to the oil tank filling pipe 56 thatleads to the tank filling port 58 that is located on or adjacent the oiltank bottom 68 of the oil tank 52.

The oil tank filling pipe 56 can take various suitable forms. It may bededicated to its function of providing a suitable passageway for oil toflow between the oil access port 55 and the tank filling port 58.However, alternatively the oil tank filling pipe 56 may be formed byusing pipe work that already exists within the engine for one or moreother purposes but is configured or is configurable to function as theoil tank filling pipe 56. This minimises the need for dedicated pipinge.g. the oil tank filling pipe 56 “Ts” into a line between the oilaccess port 55 and the oil tank 52. This assists in reducing the weightof the engine thereby increasing the efficiency of the engine andreducing specific fuel consumption (SFC).

One or more temperature and/or pressure sensors (not shown) may beprovided alongside one or more sections of the oil tank filling pipe 56and communicate with a controller (not shown) to provide an alert shouldthe oil tank filling pipe 56 burst or become blocked.

If required or desired, one or more check valves (not shown) may beprovided along the oil tank filling pipe 56 to prevent oil from flowingback through the oil pipe. However for reasons to be explained suchcheck valves are generally not required.

The tank filling port 58 can take various forms. It is located at ornear the bottom 68 of the oil tank 52 and the system relies on gravityso that when oil is poured into the oil access port 55 it flows down andthrough the oil tank filling pipe 56 and the through the oil tankfilling port 58 into the oil tank to and fill the tank as needed. Insome embodiments the tank filling port 58 is located within or adjacenta base of the oil tank 52. In the embodiment shown in FIG. 4, the tankfilling port 58 is formed in the base of the oil tank 52. However onother embodiments the tank filling port 58 is formed elsewhere in theoil tank bottom 68, for example in a side of the oil tank near the baseof the oil tank 52.

The oil tank filling system is configured so that oil that is suppliedto the oil access port 55 flows to and into the oil tank 52 usinggravitational force. This is at least in part achieved by the oil accessport 55 being located in the aft core cowl, particularly in the upperportion the engine nacelle 62 that houses the core 11 of the engine,especially approaching the top of that the engine nacelle 62, and thetank filling port 58 being located at or adjacent oil tank bottom 68 ofthe oil tank 52 of the engine. The oil tank filling pipe 56 and the tankfilling port 58 may be designed, constructed and positioned tofacilitate a smooth gravity lead passage of oil through the oil tankfilling system.

While the oil tank filling system relies on gravity to transport oilthrough the system and into the oil tank, the oil tank filling systemmay include an oil pump (not shown) that can be activated to pump oilthrough the oil tank filling pipe and the tank filling port and into theoil tank. This may reduce the time needed to fill the tank.

The oil tank filling system may be configured to circulate at least aportion of the oil around at least a part of the oil tank filling systemon a periodic or even continual basis to avoid oil stagnating within thesystem. If the gas turbine engine includes an auxiliary gear box (AGB)74 that powers a main engine oil pump, then the spill flow from thatpump may be directed via the oil tank filling system to avoidstagnation.

The oil pump can take various suitable forms and be located in varioussuitable parts of the engine to fulfil its function. In some embodimentsthe oil pump may be an auxiliary pump for an auxiliary oil system.

The oil tank filling system may include an oil pump controller (notshown) that controls the operation of the oil pump in the filling of theoil tank 52.

The oil tank filling system of the present disclosure may include asight viewing glass 76 that shows or indicates the level of oil in theoil tank 52.

The sight viewing glass 76 can take various suitable forms and belocated in various suitable parts of the engine to fulfil its function.

The sight viewing glass 76 may, for example, be located to be visiblethrough an access panel on the aft core cowl and communicate with theoil tank filling pipe 56 to visually alert when oil is backing up theoil filling tank pipe. Sighting oil backing up the oil filling tank pipeindicates the oil tank 52 is almost full or is at least approaching itspredetermined optimal volume for filling the tank rather than the totalavailable volume of the oil tank.

In some embodiments, the sight viewing glass 76 may provide visualaccess to a portion of the sight viewing glass that includes atransparent window so that a maintenance engineer can see through thesight viewing glass 76 and the transparent window when oil is backing upthe oil filling tank pipe.

In some embodiments, the sight viewing glass 76 itself provides thetransparent window into the oil tank filling pipe 56.

In some embodiments the sight viewing glass 76 is located to be visiblethrough an access panel on the aft core cowl adjacent the oil accessport 55 and communicates with the oil tank filling pipe 56 so that thelevel shown through the sight viewing glass reflects that of the oil inthe oil tank 52. In such case the sight viewing glass 76 will generallybe located on the same level (i.e. with respect to the horizon,orientated in use) as the oil access port 55. In some embodiments thesight viewing glass 76 may be located alongside the oil access port 55.In other embodiments the sight viewing glass 76 may be spaced apart fromthe oil access port 55 with suitable piping providing communicationbetween them.

In some embodiments the oil access port 55 and the sight viewing glass76 may be positioned to be outboard of a matrix cooler i.e. to takeadvantage of the cooler's presence to shield it from radiation from thecore, more specifically casings of the core.

An oil level sensor may be provided in the oil tank 52 that communicateswith a controller (not shown) in the engine to provide an alert whenthere is a predetermined volume of oil in the oil tank. Thepredetermined volume of oil may be an optimal volume for filling thetank rather than the total available volume of the oil tank. The alertmay, for example, be visual and/or aural.

A plurality of oil level sensors may be provided that communicate withthe controller to provide various alerts at predetermined points in thetank filling process. The alerts may, for example, be visual and/oraural.

One or more oil quality sensors may be provided in the oil tank 52 or inthe oil tank filling pipe 56 that measure the quality of the oil withinthe oil tank 52 or the oil tank filling pipe 56. Such oil quality sensoror sensors can take a variety of suitable forms and may communicate withan oil quality controller (not shown) that informs the maintenanceengineer and/or the pilot as to the quality of the oil currently withinthe oil tank and/or the oil tank filling system. Due to the relativeinaccessibility of the oil tank 52, an oil quantity transmitter placedin the feed pipe 56 provides in-flight indication of tank oil level in alocation which allows a more rapid replacement of the unit in the eventof a unit failure.

As mentioned above, one or more check valves (not shown) may be providedalong the oil tank filling pipe 56 to prevent oil from flowing backthrough the oil pipe. However such check valves may not be desirable forembodiments that rely on the sighting of oil backing up the oil fillingtank pipe 56 to indicate that the oil tank 52 is almost full or is atleast approaching its predetermined optimal volume for filling the tankrather than the total available volume of the oil tank.

In use, a maintenance engineer supplies oil to an aircraft on the tarmacthat requires its oil tank to be filled or simply topped up inpreparation for a flight. The maintenance engineer supplies the oil intothe oil access port 55 of the oil tank filling system 50. The oil accessport 55 is located on the aft core cowl of the engine. The oil flowsfrom the oil access port 55 through the oil tank filling pipe 56 towardsthe tank filling port 58 that is formed in the oil tank bottom 68. Theoil flows into the oil tank 52 to fill the oil tank. The oil tankfilling system is configured so that oil that is supplied to the oilaccess port 55 flows to and into the oil tank using gravitational force.

As mentioned above, the present disclosure also provides a method offilling the oil tank of a gas turbine engine that includes an aft corecowl. The method comprises the steps of: supplying oil to the oil accessport 55 located on the aft core cowl of the engine; and transporting theoil using gravitational force from the oil access port 55 through theoil tank filling pipe 56 that leads to a tank filling port 58, locatedat or adjacent the oil tank bottom 68 of the oil tank 52 located withinthe core 11 of the engine, and into the oil tank 52.

The method enables a maintenance engineer to fill an oil tank 52 towhich access is restricted, for example by the fan nacelle 21, the fancase that is covered by the fan nacelle, one or more outlet guide vanes64 or simply by being located within the core engine. The method alsoenables a generally inaccessible oil tank 52 to be filled remotely froma readily accessible oil access port 55.

The present disclosure also provides a gas turbine engine that includesthe oil tank filling system 50 of the present disclosure.

That gas turbine engine can take many suitable forms, for example thegas turbine engine may be for an aircraft wherein the gas enginecomprises an engine core comprising a turbine, a compressor, and a coreshaft connecting the turbine to the compressor; a fan located upstreamof the engine core, the fan comprising a plurality of fan blades; and agearbox that receives an input from the core shaft and outputs drive tothe fan so as to drive the fan at a lower rotational speed than the coreshaft.

The oil tank filling system of the present disclosure is especiallyuseful when the gas turbine engine is a turbofan engine that has a largefan and therefore a large fan case and a large fan nacelle thatrestricts access to the core of the engine, particularly above or belowone or more outlet guide vanes on the core engine.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

We claim:
 1. A gas turbine engine that includes a core, an aft corecowl, and an oil tank filling system, the oil tank filling systemcomprising: an oil tank that has an oil tank top and an oil tank bottomand is located within the core of the gas turbine engine; an oil accessport located on the aft core cowl, wherein oil is supplied to the oilaccess port; an oil tank filling pipe that leads from the oil accessport to a tank filling port located at or adjacent the oil tank bottom;and a sight viewing glass configured to show or indicate a level of oilin the oil tank; wherein the sight viewing glass is located on the aftcore cowl and communicates with the oil tank filling pipe to visuallyalert when oil is backing up the oil tank filling pipe; the oil tankfilling system being configured so that oil that is supplied to the oilaccess port flows to and into the oil tank using gravitational force. 2.The gas turbine engine of claim 1 wherein the oil tank is shaped tocurve around the gas turbine engine within the core of the gas turbineengine.
 3. The gas turbine engine of claim 1 further comprising outletguide vanes, wherein the oil tank is located between opposing ones ofthe outlet guide vanes.
 4. The gas turbine engine of claim 1 wherein thesight viewing glass is located to be visible through an access panel onthe aft core cowl adjacent the oil access port.
 5. The gas turbineengine of claim 1 wherein the oil access port and the sight viewingglass are positioned to be outboard of a matrix cooler.
 6. The gasturbine engine of claim 1 further comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; a fan locatedupstream of the core, the fan comprising a plurality of fan blades and agearbox that receives an input from the core shaft and outputs drive tothe fan so as to drive the fan at a lower rotational speed than the coreshaft, wherein the gas turbine engine is for use on an aircraft.
 7. Amethod of filling an oil tank of a gas turbine engine that includes acore and an aft core cowl, the method comprising the steps of: supplyingoil to an oil access port located on the aft core cowl of the gasturbine engine; transporting the oil using gravitational force from theoil access port through an oil tank filling pipe that leads to a tankfilling port, located at or adjacent an oil tank bottom of the oil tanklocated within the core of the gas turbine engine, and into the oiltank; and providing a sight viewing glass configured to show or indicatea level of oil in the oil tank, wherein the sight viewing glass islocated on the aft core cowl and communicates with the oil tank fillingpipe to visually alert when oil is backing up the oil tank filling pipe.